This invention relates generally to modular spacecraft engineering and, more particularly, to spacecraft electronic equipment modules that can be externally mounted on a core spacecraft structure. Conventional spacecraft designs have placed electronic equipment on panels located inside box-like modules, with the panels typically forming the walls of the modules. The modules are mounted onto a spacecraft and interconnected with complex interconnecting wiring. Heat dissipated from the electronics equipment is conducted into the module panels and radiates into space. The size of each module is determined in part by the heat radiating area needed to cool the equipment, so there is often unused volume within each module box or compartment. Because only one side of the equipment panel is used as a radiator, the modules often have to be very large to maintain desired operating temperatures. A related problem is that accessing equipment requires the removal of panels from the equipment compartment. Not only are modules of this type bulky to accommodate in a launch vehicle and difficult to access for servicing, but they typically need to be coupled to other thermal radiator panels, through heat-conducting pipes, to provide overall thermal management of the spacecraft.
Further, because the box modules are rigidly bolted to the core structure of the spacecraft, thermally induced stresses are a significant problem because of temperature differences between the modules and the spacecraft core structure. In brief, these conventional equipment modules are structurally and thermally dependent on the spacecraft core structure, and the overall design of the spacecraft must take into account the thermal requirements of each module and the structural forces resulting from the presence of each module.
As a result of these difficulties, the spacecraft core structure is usually constructed to have a relatively high weight and volume, to support the modules and to provide an adequate thermal radiation area. Also, it is usually the case that the choice of materials of both the modules and the spacecraft core structure is limited because there is a concern for differential thermal expansion. A further difficulty is that removal or addition of a module upsets the overall structural and thermal design to some degree. Alignment problems, thermal management problems, or both, can result from simply removing or adding a module.
All of the foregoing problems are attributable to interdependence of the modules and the spacecraft core structure, which together interact, both structurally and thermally, as parts of a larger assembly. It will be appreciated, therefore, that there is a need for a different approach to the construction of spacecraft modules for supporting electronic equipment, to overcome the difficulties noted above. The present invention is directed to the undesirable thermal dependence of modules on the spacecraft as whole.